Axially-split radial turbines

ABSTRACT

Embodiments of an axially-split radial turbine, as are embodiments of a method for manufacturing an axially-split radial turbine. In one embodiment, the method includes the steps of joining a forward bladed ring to a forward disk to produce a forward turbine rotor, fabricating an aft turbine rotor, and disposing the forward turbine rotor and the aft turbine rotor in an axially-abutting, rotationally-fixed relationship to produce the axially-split radial turbine.

CROSS-REFERENCE TO RELATED APPLICATION

This application a divisional of co-pending U.S. application Ser. No.13/444,070, filed Apr. 11, 2012.

TECHNICAL FIELD

The present invention relates generally to gas turbine engines and, moreparticularly, to axially-split radial turbines for usage within gasturbine engines, as well as to methods for fabricating axially-splitradial turbines.

BACKGROUND

Gas turbine engines (“GTE”) have been engineered extensively to improveefficiency, thrust-to-weight ratios, and other measures of engineperformance. One of the most direct manners by which engine performancecan be improved is through increases in core rotational speeds andturbine inlet temperatures generated during engine operation. However,as turbine inlet temperatures and rotational speeds increase, so too dothe thermal and mechanical demands placed on the GTE components. Themost demanding performance requirements are typically placed on the highpressure turbine rotor or rotors, which are positioned immediatelydownstream of the GTE combustion section and which rotate at thegreatest speed during engine operation. The turbine blades, inparticular, are directly exposed to combustive gas flow at or near peaktemperatures and are consequently heated to exceedingly hightemperatures at which most alloys weaken or melt and become prone tooxidation or other forms of chemical degradation. By comparison, theinner annular body of the turbine (commonly referred to as the “turbinedisk”) is largely shielded from direct exposure to combustive gas flow,but is subject to considerable mechanical stress resulting from bothcentrifugal forces acting on the turbine rotor at high rotational speedsand rim-to-bore thermal gradients.

Turbines can be broadly divided into two categories, axial and radialturbines, based upon the direction of airflow received by the turbinerelative to the turbine's rotational axis. Compared to axial turbines,radial turbines offer certain performance benefits including superiorpressure ratios; e.g., a single radial turbine is often capable ofproviding a power output equivalent to two similarly-sized axialturbines. Radial turbines also provide relatively smooth exit flowconditions as compared to axial turbine turbines, which are oftencharacterized by highly turbulent outflow. However, by inherent design,radial turbines tend to be relatively lengthy in an axial direction and,thus, have lower power-to-weight ratios. In addition, it is difficult tofabricate radial turbines having internal cooling features, such ascooling flow passages formed within the turbine blades. Specifically,casting of radial wheels with internally-cooled blades has provenexcessively costly due to low yields results from the usage of complexcasting cores generally required to form the blade cooling passages. Anuncooled radial turbine may be incapable of withstanding prolongedstresses at high operational speeds without premature fatigue insituations wherein a gas turbine engine is operated at relatively highcombustive gas flow temperatures to boost power output. In suchsituations, the inability to manufacture an internally-cooled radialturbine in a reliable and cost-effective manner may prevent the usage ofa radial turbine altogether and a number of cooled axial turbines may beemployed instead. For at least these reasons, radial turbines tend to beutilized only within a relatively limited number of small, low-flow gasturbine engine platforms.

It is thus desirable to provide embodiments of a radial turbine suitablefor usage in a gas turbine engine that is relatively lightweight andcost effective to implement. Ideally, embodiments of such a radialturbine would include internal cooling passages to permit operation ofthe radial turbine at higher temperatures. It would also be desirableif, in certain embodiments, the different sections of such a radialturbine were fabricated from disparate alloys tailored to the particularoperating conditions experienced by each turbine section. Finally, itwould be desirable to provide embodiments of a method for fabricatingsuch a radial turbine. Other desirable features and characteristics ofembodiments of the present invention will become apparent from thesubsequent Detailed Description and the appended Claims, taken inconjunction with the accompanying drawings and the foregoing Background.

BRIEF SUMMARY

Embodiments of a method for producing an axially-split radial turbineare provided. In one embodiment, the method includes the steps ofjoining a forward bladed ring to a forward disk to produce a forwardturbine rotor, fabricating an aft turbine rotor, and disposing theforward turbine rotor and the aft turbine rotor in an axially-abutting,rotationally-fixed relationship to produce the axially-split radialturbine.

Embodiments of an axially-split radial turbine are further provided. Inone embodiment, the axially-split radial turbine includes a forwardturbine rotor and an aft turbine rotor, which is disposed axiallyadjacent to and rotationally fixed relative to the forward turbinerotor. The forward rotor includes a forward disk and a forward bladedring, which circumscribes and is metallurgically bonded to the forwarddisk.

BRIEF DESCRIPTION OF THE DRAWINGS

At least one example of the present invention will hereinafter bedescribed in conjunction with the following figures, wherein likenumerals denote like elements, and:

FIG. 1 is a simplified schematic of an exemplary gas turbine engine(partially shown) including an internally-cooled, axially-split radialturbine, as illustrated in accordance with an exemplary embodiment ofthe present invention;

FIG. 2 is an isometric view of the axially-split radial turbine shown inFIG. 1;

FIG. 3 is a cross-sectional view of a portion of the axially-splitradial turbine shown in FIGS. 1 and 2, as taken along line 3-3identified in FIG. 2, and depicting the fore and aft blade segments in aflattened, cutaway view to more clearly illustrate exemplary internalcooling features, which can be formed within the blade segments;

FIG. 4 is a flowchart illustrating an exemplary method for manufacturingan axially-split radial turbine, such as the internally-cooled,axially-split radial turbine shown in FIGS. 1-3;

FIG. 5 is an isometric view of one of a plurality of forward bladedpieces that may be produced and subsequently inter-bonded with aplurality of like blade pieces to produce the forward bladed ring of theaxially-split radial turbine, in accordance with the manufacturingmethod shown in FIG. 4;

FIG. 6 is an isometric view of a bladed ring formation into which theexemplary bladed piece shown in FIG. 5 and like bladed pieces may bearranged prior to consolidation into a bonded bladed ring;

FIGS. 7 and 8 are exploded and assembled views, respectively, of athermal growth constraint tool suitable for carrying-out a directedthermal growth bonding process to consolidate the bladed ring formationshown in FIG. 6 and produce a monolithic bladed ring;

FIG. 9 is a cross-sectional view of the thermal growth constraint tooland ring formation taken along line 9-9 identified in FIG. 8;

FIG. 10 schematically illustrates the thermal growth constraint tool andring formation shown in FIGS. 7-9 prior to metallurgical consolidationinto a monolithic or single-piece ring;

FIGS. 11-13 are detailed views of a portion of the thermal growthconstraint tool and ring formation shown in FIG. 10 illustrating onemanner in which the outward radial growth of the bladed ring formationmay be impeded by the constraint tool during a directed thermal growthbonding process to promote shank-to-shank bonding of the forward bladedpieces; and

FIGS. 14 and 15 are isometric views of a partially-machined disk forging(in this example, the forward disk) prior to and after joinder to themachined bonded bladed ring (in this example, the forwarded bladed ring)to yield the forward turbine rotor.

DETAILED DESCRIPTION

The following Detailed Description is merely exemplary in nature and isnot intended to limit the invention or the application and uses of theinvention. Furthermore, there is no intention to be bound by any theorypresented in the preceding Background or the following detaileddescription.

FIG. 1 is a simplified cross-sectional view of a portion of a gasturbine engine (GTE) 18 including an internally-cooled, axially-splitradial turbine 20, as illustrated in accordance with an exemplaryembodiment of the present invention. The term “axially-split radialturbine,” as appearing herein, is utilized in a broad sense to denote aradial turbine that includes or that is fabricated from at least twoaxially-mating or axially-abutting structures or sections. Asillustrated in FIG. 1 and described herein, GTE 18 is offered by way ofexample only to provide a convenient and non-limiting context in whichan exemplary embodiment of radial turbine 20 can be readily understood.It will be readily appreciated that embodiments of axially-split radialturbine 20 can be employed within various different types of gas turbineengine including, but not limited to, other types of turbofan,turboprop, turboshaft, and turbojet engines, whether deployed onboard anaircraft, watercraft, or ground vehicle (e.g., a tank), included withinan auxiliary power unit, included within industrial power generators, orutilized within another platform or application.

In the exemplary embodiment shown in FIG. 1, GTE 18 is illustrated as atwo spool engine including a low pressure (LP) shaft 22 and highpressure (HP) shaft 24. LP shaft 22 and HP shaft 24 are co-axial; thatis, LP shaft 22 extends through a longitudinal channel provided throughHP shaft 24 along the engine centerline and the rotational axis of LPshaft 22, HP shaft 24, and radial turbine 20 (represented in FIG. 1 bydashed line 26). GTE 18 further includes a compressor section 28, acombustion section 30, and a turbine section 32, each of which is onlypartially shown in FIG. 1. An HP compressor 33 is included withincompressor section 28 and mounted to the upstream or forward end of HPshaft 24. Axially-split radial turbine 20 is contained within turbinesection 32 and mounted to the opposing, aft end of HP shaft 24. Althoughnot shown in FIG. 1 for clarity, compressor section 28 and turbinesection 32 may further include an LP compressor and an LP turbine,respectively, which may be mounted to opposing ends of LP shaft 22.Alternatively, in embodiments wherein GTE 18 is a turbofan-type engine,an intake fan may be mounted to the upstream end of LP shaft 22.

During engine operation, airflow is supplied to compressor section 28 byan intake section included within GTE 18 (not shown). A serpentinecompressor section duct 36 guides the pressurized airflow through HPcompressor 33, which rotates to further compress the airflow prior todischarge into combustion section 30 (represented in FIG. 1 by arrows34). Within combustion section 30, the compressed airflow is mixed withfuel and ignited to produce combustive gasses. As indicated in FIG. 1 byarrows 38, the combustive gasses expand rapidly, exit combustion section30 through a turbine scroll 40, and are directed through HP turbine 20.The combustive gas flow drives rotation of HP turbine 20, which drivesfurther rotation of HP shaft 24 and HP compressor 33. After flowingthrough HP turbine 20, the combustive gases flow through a turbinesection duct 42 and into a turbine nozzle ring 44 (indicated by arrows46). Turbine nozzle ring 44 accelerates the combustive gas flow andimparts the gas flow with a desired tangential component prior toreaching an LP turbine (not shown), which is positioned downstream ofnozzle ring 44 and mounted to the aft end of LP shaft 22. The combustivegas flow drives the rotation of the non-illustrated LP turbine, whichdrives further rotation of LP shaft 22. The rotation of shafts 22 and 24provides power output, which may be utilized in a variety of differentmanners depending upon whether GTE 18 assumes the form of a turbofan,turboprop, turboshaft, turbojet engine, or an auxiliary power unit, tolist but a few examples. The combustive gas flow is then mixed withcooler bypass flow and exhausted from GTE 18; e.g., in the case ofturbofan or turbojet engine, the combustive gas flow may be exhaustedthrough a propulsion nozzle to provide forward thrust.

Advantageously, internally-cooled, axially-split radial turbine 20increases the overall efficiency of GTE 18 by allowing increasedcombustion temperatures and higher efficiency cycles. The improvedthermal and mechanical capabilities of radial turbine 20 are achieved,at least in part, by fabricating the turbine blades to include internalcooling flow passages (hidden from view in FIG. 1). As noted in theforegoing section entitled “BACKGROUND,” casting of radial wheels withcooled blades has traditionally been undesirably costly due toexcessively low yields as a result of complex casting cores required forthe blade cooling passages. To overcome this limitation, the followingdescribes exemplary embodiments of a method for fabricating aninternally-cooled radial turbine, such as radial turbine 20, wherein aplurality of individual bladed pieces are first cast and subsequentlymetallurgically consolidated into a monolithic bladed ring, which isthen bonded onto a forward disk to yield a forward turbine rotor. Asimilar process may then be carried-out to produce an aft turbine rotor,which is then placed in an axially-abutting relationship with theforward turbine rotor to produce the radial turbine. In addition, such aprocess allows the bladed pieces (and thus the forward and/or aftsections of the turbine blades) to be cast from single crystalsuperalloy materials having optimal resistances to oxidation and otherforms of chemical degradation and good mechanical strength at hightemperature operating conditions, while other portions of the radialturbine (e.g., the forward and aft disks) are fabricated from non-singlecrystal alloys having relatively high stress rupture and low cyclefatigue properties. An exemplary embodiment of the axially-split radialturbine fabrication method is described below in conjunction with FIGS.4-15; first, however, the structure of radial turbine 20, in anillustrative, but non-limiting implementation, is described in detailbelow in conjunction with FIGS. 2 and 3.

FIG. 2 is an isometric view of internally-cooled, axially-split radialturbine 20 employed with GTE 18 (FIG. 1). Referring collectively toFIGS. 1 and 2, axially-split radial turbine 20 is comprised of at leasttwo discrete parts, which are matingly assembled to yield the completedturbine. In the illustrated example, specifically, axially-split radialturbine 20 is comprised of two mating halves, namely, a forward turbinerotor 50 and an aft turbine rotor 52. When radial turbine 20 isassembled, the aft face of forward turbine rotor 50 and the forward faceof aft turbine rotor 52 abut along an annular interface, which resideswithin a plane generally orthogonal to rotational axis 26 and which isreferred to herein as “split line 54.” Forward turbine rotor 50 and aftturbine rotor 52 can be metallurgically bonded utilizing, for example, asoldering, brazing, or diffusion bonding technique. It is preferred,however, that forward turbine rotor 50 and aft turbine rotor 52 arediscrete, non-bonded parts, which are maintained in intimate contact orengagement utilizing a mechanical means, such as a plurality offasteners or an axial clamping device. For example, as shown in FIG. 1,a curvic may exert an axial clamping preload through forward turbinerotor 50 and aft turbine rotor 52 maintaining rotors 50 and 52 in anabutting, rotationally-fixed engagement.

FIG. 3 is a cross-sectional view of a portion of radial turbine 20, astaken along line 3-3 in FIG. 2. As can be seen in FIGS. 2 and 3, forwardturbine rotor 50 includes a forward hub section 56 and a plurality offorward blade segments 58, which are circumferentially spaced around andextend radially outward from hub section 56. Aft turbine rotor 52likewise includes an aft hub section 60 and a plurality of aft bladesegments 62, which are circumferentially spaced around and extendradially outward from hub section 60. As shown most clearly in FIG. 2,forward blade segments 58 and aft blade segments 62 wrap or twisttangentially about the rotational axis of radial turbine 20 in adirection substantially opposite the direction of rotation. However,forward blade segments 58 and aft blade segments 62 are depicted in aflattened-out, cutaway view in FIG. 3 to more clearly illustrate variousinternal cooling features 64, 66 that can be formed in forward bladesegments 58 and/or aft blade segments 62 pursuant to the below-describedfabrication process. Internal cooling features 64, 66 may include anynumber and combination of internal cooling channels, slots, apertures,and the like suitable for cooling blade segments 58 and 62 andneighboring regions of radial turbine 20 during high temperatureoperation; e.g., as indicated in FIG. 3, the internal cooling features64, 66 may include blade cooling passages 64 formed in forward bladesegments 58 (only one of which is shown in FIG. 3) and/or blade coolingpassages 66 formed in aft blade segments 62 (again, only one of which isshown).

Forward hub section 56 aligns axially with aft hub section 60 and,preferably, radially pilots thereto. Although not illustrated in FIGS. 2and 3, a radially-overlapping hub interface (e.g., an annular lap joint)may be provided between forward hub section 56 and aft hub section 60for radial piloting purposes. Furthermore, as shown in FIG. 3, a centralopening 65 provided through forward hub section 56 aligns with a centralopening 67 provided through aft hub section 60 to define a longitudinalchannel 65, 67 through radial turbine 20. A spool (e.g., high pressureshaft 24 of GTE 18) extends through this longitudinal channel whenradial turbine 20 is installed within GTE 18 (FIG. 1). Radial turbine 20may be fixedly mounted to the gas turbine engine shaft utilizing, forexample, a curvic-type attachment. Forward blade segments 58 alignaxially with and preferably abut or interlock with aft blade segments62. Axial alignment of forward blade segments 58 and aft blade segments62 may be maintained by an axial clamping force or pre-load exerted onradial turbine 20 by a tie-shaft (not shown). Additionally, the matinginterface between forward hub section 56 and aft hub section 60 may befabricated to include one or more alignment features (e.g., keys, teeth,or castellations). As a result of this structural configuration,neighboring pairs of forward blade segments 58 and aft blade segments 62cooperate to form continuous blade structures 58, 62, which, along withouter hub flow surfaces 69 (FIG. 3), provide substantially uninterruptedairflow guidance surfaces when transitioning from forward hub section 56and aft hub section 60 of radial turbine 20.

By virtue of its multi-piece construction, axially-split radial turbine20 can be readily fabricated to include at least one inner disk cavityfor mass reduction and/or for internal cooling purposes. For example, asshown in FIG. 3, axially-split radial turbine 20 may be fabricated toinclude an inner disk cavity 68, 70 formed within the body of turbine 20between forward hub section 56 and aft hub section 60. In theillustrated example, inner disk cavity 68, 70 assumes the form of anannular void extending around the rotational axis of radial turbine 20;however, in further embodiments, inner disk cavity 68, 70 may comprise aplurality of discrete cavities or pockets, which are formed within thebody of turbine 20 as an angularly- or circumferentially-spaced array.Inner disk cavity 68, 70 is conveniently formed in trailing face offorward hub section 56, the leading face of aft hub section 60, or acombination of the two by removal of material during the below-describedfabrication process. For example, a first annular channel or groove 68may be formed in the trailing face of forward hub section 56 and asecond annular channel or groove 70 may be formed in the leading face ofaft hub section 60 by machining, which then align to form inner diskcavity 68, 70 when radial turbine 20 is assembled.

In addition to providing the above-noted benefits in mass reduction,inner disk cavity 68, 70 is also advantageously utilized to enhance theinternal cooling performance of radial turbine 20. To this end,axially-split radial turbine 20 can be fabricated to further include aplurality of hub cooling passages 72 (one of which is shown in FIG. 3),which place inner disk cavity 68, 70 in fluid communication with acooling air source. As indicated in FIG. 3, hub cooling passages 72 areconveniently formed as radially-extending conduits, which arecircumferentially spaced about the rotational axis of turbine 20 andwhich extend along the interface between forward hub section 56 and afthub section 60. Each hub cooling passage 72 includes an inlet 76, whichis formed in the inner circumferential surface of turbine hub 56, 60 inradial alignment with split line 54; and an outlet 74, which is fluidlycoupled to inner disk cavity 68, 70. Inner disk cavity 68, 70 is, inturn, fluidly coupled to the inlet or inlets of one or more cooling flowpassages formed within forward blade segments 58 and/or aft bladesegments 62. For example, as shown in FIG. 3, inner disk cavity 68, 70may be fluidly coupled to an inlet 75 of a first flow passage 64 formedwithin each forward blade segment 58 and to an inlet 77 of a second flowpassage 66 formed within each aft blade segment 62. In this manner,cooling airflow received along inner channel 65, 67 can be directed intoinner disk cavity 68, 70 by hub flow passages 72 and, from cavity 68,70, distributed to blade cooling passages 64, 66.

Describing an exemplary radial turbine cooling flow scheme, andreturning briefly to FIG. 1, relatively cool airflow may be bled from asection of the engine upstream of combustion section 30 and supplied toradial turbine 20 during engine operation; e.g., as indicated in FIG. 1by arrows 80, a portion of the airflow flowing within compressor sectionduct 32 may be diverted into outlet 82 of a combustion bypass conduit84. A portion of this relatively cool bypass airflow may be directedtoward the leading edge of forward hub section 56 to provide saddlecooling (indicated in FIG. 1 by arrow 86), while the remainder of thebypass air may be directed radially inward toward the engine centerline.The bypass airflow may then be conducted axially along HP shaft 24 andultimately flow into inlets 76 of hub cooling passages 72. As indicatedin FIG. 1 by arrows 88, hub cooling passages 72 direct the relativelycool bypass airflow radially outward through the body of turbine 20 andinto inner disk cavity 68, 70, which then distributes the coolingairflow to internal cooling features 64, 66 provided in the turbineblade segments (FIG. 3). Various different devices and techniques canalso be utilized to encourage cooling airflow through theabove-described cooling circuit. For example, radial turbine 20 can beequipped with one or more spoiler tubes. Additionally or alternatively,GTE 18 may be equipped with a “Tangential On-Board Injection” or “TOBI”device 90 (FIG. 1), which injects the cooling airflow immediatelyupstream of the high pressure turbine while imparting the airflow with atangential or swirling-type motion to reduce parasitic pumping losses.

As previously stated, forward turbine rotor 50 is advantageouslyproduced from two distinct components or sections, which aremetallurgically bonded together during fabrication to yield a monolithicor unitary piece. In the exemplary embodiment illustrated in FIGS. 2 and3, and referring specifically to FIG. 3, the two distinct components orsections bonded together to produce forward turbine rotor 50 are aforward bladed ring 94 and a forward disk 96. Forward bladed ring 94includes or embodies a majority of each forward blade segment 58 and, incertain cases, may include the entirety of each forward bladed segment58. By comparison, forward disk 96 includes or embodies the entirety orsubstantial entirety of forward hub section 56. The demarcation betweenforward bladed ring 94 and forward disk 96 is visually indicated in FIG.3 by dashed bond line 92. As can be seen, dashed bond line 92 has agenerally conical geometry, which decreases in outer diameter whenmoving in a fore-aft direction. By comparing bond line 92 to theillustrated outer hub flow surface 69, it can further be seen that aleading portion of the hub flow path is defined by forward disk 96,while an intermediate portion of the hub flow path is defined by forwardbladed ring 94. Inner disk cavity 68, 70 is bounded, in part, by bothforward bladed ring 94 and forward disk 96. As forward bladed ring 94extends to and partially defines inner disk cavity 68, 70, blade coolingpassages 64 can be formed entirely within forward bladed ring 94, whichcan be readily fabricated from a number of small-volume cast pieces tofacilitate the formation of blade cooling passages 64, as described morefully below.

As is forward turbine rotor 50, aft turbine rotor 52 can be fabricatedfrom two distinct components or sections, which are metallurgicallybonded together during fabrication to yield a monolithic or unitarypiece. Referring to FIG. 3, an aft bladed ring 100 and an aft disk 102are bonded together to yield aft turbine rotor 52. Aft bladed ring 94includes or embodies a majority of each aft blade segment 62 and, in theillustrated example, includes the entirety of each aft blade segment 62and a relatively thin annular region of aft hub section 60 underlyingblade segments 62, as taken in an inward radial direction. As acorollary, aft bladed ring 94 defines the aft sections of hub flow paths69. The demarcation between aft bladed ring 100 and aft disk 102, whichincludes or embodies the entirety or substantial entirety of aft hubsection 60, is visually indicated in FIG. 3 by dashed bond line 98. Inaddition to being bound by forward bladed ring 94 and forward disk 96,inner disk cavity 68, 70 is further bound by aft bladed ring 100 and aftdisk 102. In this manner, blade cooling passages 66 can be formedentirely within aft bladed ring 100, which can be readily fabricatedfrom a number of small-volume cast pieces to facilitate the formation ofpassages 66. The foregoing notwithstanding, aft turbine rotor 52 can befabricated from a single piece (e.g., a unitary forging or single-shotcasting) in alternative embodiments and thus may not include bond line98.

FIG. 4 is a flowchart illustrating a method 110 for manufacturing anaxially-split radial turbine in accordance with an exemplary embodimentof the present invention. For ease of explanation, the following willdescribe exemplary method 110 in conjunction with the fabrication ofinternally-cooled, axially-split radial turbine 20 shown in FIGS. 1-3;however, method 110 can be utilized to produce various other types ofradial turbines having structural features that differ from thosedescribed above-in conjunction with radial turbine 20. Several of thesteps described in conjunction with method 110 below are similar tosteps in the fabrication processes described in co-pending U.S. patentapplication Ser. No. 13/371,922, filed with the USPTO on Feb. 13, 2012,and assigned to the assignee of the instant application.

Forward turbine rotor 50 and aft turbine rotor 52 can be produced inparallel or in series, with either rotor produced first. In exemplarymethod 110, forward turbine rotor 50 is initially produced by way of afirst sub-process (SUB-PROCESS BLOCK 112, FIG. 4). To commence theproduction of forward turbine rotor 50, a plurality of bladed pieces isproduced or otherwise obtained (STEP 114, FIG. 4). The bladed pieces maybe purchased from a supplier or independently fabricated by, forexample, individual casting and machining. The bladed pieces are ideallyfabricated from an oxidation-resistant material having good mechanicalstrength at high temperatures. In a preferred embodiment, each bladedpiece is cast from a single crystal superalloy having itscrystallographic grain structure oriented to provide optimal mechanicalstrength in a radial direction. By way of non-limiting example, anoxidation-resistant, single crystal, nickel-based superalloy may beutilized, such as the nickel-based superalloy commercially identified as“CMSX 4” or the nickel-based superalloy commercially identified as“SC180,” although various other high temperature materials may also beutilized including equiaxed superalloys. Notably, by initially casting anumber of small volume bladed pieces, casting yields can be favorablyincreased as compared to the casting of a bladed ring in its entirety.In addition, casting such small-volume bladed pieces allows the turbineblades to be imparted with relatively complex internal geometries and,therefore, relatively tortuous internal cooling flow passages, such asblade cooling passages 64 and 66 shown in FIG. 3.

FIG. 5 is an isometric view of an exemplary bladed piece 116 that may beproduced pursuant to STEP 114 of method 110 (FIG. 4). In this particularexample, bladed piece 116 includes a wedge-shaped body or shank 118 fromwhich a forward blade segment 58 extends. Shank 118 has been cast orotherwise fabricated to an outer sacrificial portion (e.g., extra axialstock) enveloping or encompassing the desired shape of the completedshank. Stated differently, shank 118 is a rough casting or preformphysically embodying the finished shank, which may be subsequentlydefined by further machining. Shank 118 includes a sacrificial forwardshoulder 122 and a sacrificial aft shoulder 124, which extend from thecentral body of shank 118 in opposing axial directions. Opposing axialshoulders 122 and 124 are fabricated to include outer tooling contactsurfaces 126 and 128, respectively, which are located axially adjacentto forward and aft ends of forward blade segment 58 and which have agently curved or arced outer geometry. Shank 118 also includes opposinglateral faces or sidewall surfaces 130, only one of which can be seen inFIG. 5. Sidewall surfaces 130 are each angled to impart shank 118 with awedge-shaped geometry, when viewed from the forward/leading oraft/trailing end thereof. As each sidewall surface 130 of shank 118 willultimately be bonded to a mating sidewall surface of a neighboringshank, sidewall surfaces 130 are also referred to herein as“shank-to-shank bonding surfaces 130.” As described more fully below inconjunction with FIGS. 11-13, in certain embodiments, one or more stripsof braze foil may be inserted between neighboring shanks 118 prior tothe below-described bonding process. To accommodate the braze foil, eachbladed piece 116 may be fabricated to include at least one braze foilpocket 129, which may be formed in one or both of sidewall surfaces 130.

A series of steps is next performed to inter-bond or consolidate bladedpieces 116 and produce the forward bladed ring. First, bladed pieces 116are arranged in an annular group or ring formation (STEP 131, FIG. 4),such as ring formation 132 shown in FIG. 6. When bladed pieces 116 arearranged in this manner, each shank 118 is positioned between twoneighboring shanks 118 (three of which are labeled in FIG. 6), and eachshank-to-shank bonding surface 130 (identified in FIG. 5) is positionedadjacent a neighboring bonding surface 130. The neighboring bondingsurfaces 128 may contact each other at this stage in the fabricationprocess or may instead be separated by a small radial gap. Outer toolingcontact surfaces 126 and 128 of axially-extending shoulders 122 and 124(FIG. 5) combine to impart ring formation 132 with forward and aftcomposite tooling contact surfaces 134 and 136, respectively. Compositetooling contacts surfaces 134 and 136 each have a substantially annulargeometry and extend from circumferentially-spaced blade segments 58 ofring formation 132 in opposing axial directions. Due to theinwardly-tapered conical geometry of ring formation 132, forwardcomposite tooling contact surface 134 has a greater outer diameter thandoes aft composite tooling contact surface 136.

In certain implementations of method 110, bonding material may bedisposed between neighboring shank-to-shank bonding surfaces 130 (FIG.5) during STEP 131 of exemplary method 110 (FIG. 4). In embodimentswherein braze foil pockets are formed in shank-to-shank bonding surfaces130, such as pocket 129 formed in sidewall 130 (FIG. 5), a strip ofbraze foil may be inserted into each pocket. In this case, it ispreferred that the volume of each braze foil sheet is slightly greaterthan the volume of its corresponding pocket to ensure the creation of avoid-free, zero-defect braze joint during the below-described directedthermal growth bonding process. In this regard, the braze foil sheet maybe trimmed to have planform dimensions (e.g., a height and width)substantially equivalent to the dimensions of braze foil pockets, whilethe thickness of the braze foil sheet may be chosen to be slightlygreater than the pocket depth; e.g., in one implementation wherein thepocket depth is about 0.0025 inch, the braze foil thickness may be about0.0030 inch. Additionally or alternatively, braze paste may be dispensedinto pockets 129 (FIG. 5) or otherwise applied between theshank-to-shank bonding surfaces 130 (FIG. 5) to ensure a sufficientquantity of braze material is present to reliable create a void-free,zero-defect braze joint.

Next, at STEP 138 (FIG. 4), ring formation 132 is subjected to adirected thermal growth bonding process to inter-bond or arch-bondbladed pieces 116 and thereby yield a monolithic bladed ring. During thedirected thermal growth bonding process, ring formation 142 is heated apredetermined bonding temperature or temperatures while the outwardradial growth of ring formation is constrained 142 to urge convergentgrowth of the shank-to-shank bonding surfaces 130 (FIG. 5) and therebypromote shank-to-shank bonding. A specialized thermal growth constrainttool is utilized to constrain or physically impede the outward radialgrowth of ring formation 132 during the directed thermal growth process,while avoiding bending, distortion, or warping of turbine blade segments58. The thermal growth constraint tool thus enables the inter-bonding orarch-bonding of the forward bladed pieces to form a bonded bladed ring,which can then be bonded onto a forward rotor disk to produce a forwardturbine rotor, such as forward turbine rotor 50 shown in FIGS. 1-3. Anexample of a thermal growth constraint tool suitable for this purposesis described below in conjunction with FIGS. 7-9.

FIGS. 7 and 8 are exploded and assembled views, respectively, of athermal growth constraint tool 140 suitable for usage in themetallurgical consolidation of ring formation 132. Thermal growthconstraint tool 140 and ring formation 132 are also partially shown inFIG. 9, which is a cross-sectional view taken along line 9-9 identifiedin FIG. 8. In the exemplary embodiment shown in FIGS. 7-9. thermalgrowth constraint tool 140 includes four main components: (i) firstthermal growth constraint ring 142, (ii) a second thermal growthconstraint ring 144, (iii) a plurality of spacer members 146, and (iv) abaseplate 148. Thermal growth constraint rings 142 and 144, spacermembers 146, and baseplate 148 are sized and shaped to be assembledaround ring formation 132 in a stacked formation. That is, when thermalgrowth constraint tool 140 is fully assembled, thermal growth constraintring 144 is positioned over and supported by baseplate 148, spacermembers 146 are positioned over and extend upward from constraint ring144, and constraint ring 142 is positioned over and supported by spacermembers 146. To retain constraint rings 142 and 144 and spacer members146 in their desired positions after tooling assembly, a plurality ofretainers 150 (e.g., L-shaped brackets) are affixed to an outerperipheral flange 152 provided around baseplate 148 utilizing aplurality of bolts 154 or other such fasteners (shown in FIGS. 8 and 9).

Thermal growth constraint rings 142 and 144 are shaped and sized toextend around different circumferential portions of ring formation 132when loaded into thermal growth constraint tool 140. In particular,thermal growth constraint rings 142 and 144 are sized and shaped to bedisposed around composite tooling contact surfaces 134 and 136,respectively; e.g., the inner diameters of the central openings providedin rings 142 and 144 are fabricated to be slightly larger than the outerdiameters of contact surface 134 and 136, respectively. As the outerdiameter of composite tooling contact surface 134 is greater than theouter diameter of composite tooling contact surface 136, the innerdiameter of central opening provided in thermal growth constraint ring142 is likewise chosen to be greater than the inner diameter of thecentral opening provided in thermal growth constraint ring 144.

When thermal growth constraint tool 140 is assembled, thermal growthconstraint rings 142 and 144 are axially spaced or offset by spacermembers 146 such that a relatively large axial gap or annular void 156(identified in FIG. 9) is provided between the inner faces of constraintrings 142 and 144. As shown in FIGS. 8 and 9, forward blade segments 58extend radially into an annular void 156 (FIG. 9) when ring formation132 is positioned within thermal growth constraint tool 140 such thatturbine blade segments 58 generally reside between inner radial faces ofconstraint rings 142 and 144. Spacer members 146 may assume any formsuitable for maintaining the axial spacing of thermal growth constraintrings 142 and 144 during thermal processing. In the illustrated example,spacer members 146 comprise three stand-off pins, which arecircumferentially spaced around the central axis of thermal growthconstraint tool 140 and which extend from constraint ring 142 tobaseplate 148. As shown in FIG. 9, the opposing ends of each stand-offpin may be received within axially-aligning cylindrical recesses 158provided in the inner radial faces of constraint ring 142 and baseplate148.

Thermal growth constraint rings 142 and 144 are fabricated to grow in aradially outward direction at a rate significantly slower than does ringformation 132 during thermal processing. To ensure a substantiallyuniform application of force to the forward and aft ends of ringformation 132, constraint rings 142 and 144 are also preferablyfabricated to grow radially outward at substantially equivalent rates.To this end, constraint rings 142 and 144 may each be fabricated from arefractory material having a coefficient of thermal expansion (“CTE”)less than that of ring formation 132 and, preferably, less thanapproximately one half the CTE of the material from which the individualbladed pieces are formed. It is also desirable for constraint rings 142and 144 to have a sufficient high temperature strength to constrain theoutward radial growth of ring formation 132, as described more fullybelow. Considering this, it is preferred that thermal growth constraintrings 142 and 144 are formed from a metallic material (i.e., anon-ceramic metal or superalloy) having a relatively low CTE, such asmolybdenum or rhenium. In one embodiment, thermal growth constraintrings 142 and 144 each comprise an annular body fabricated frommolybdenum.

Spacer members 146 are preferably fabricated from a material having CTEsubstantially equivalent to (e.g., differing by no more than about 20%relative to) the CTE of the bladed pieces from which ring formation 132is formed. When formed from a material having a CTE similar to the CTEof bladed pieces, spacer members 146 grow in axial direction at a ratesubstantially equivalent to ring formation 132 to maintain proper radialalignment between thermal growth constraint rings 142 and 144 andtooling contact surfaces 134 and 136, respectively. This also helps tomaintaining proper axial spacing between constraint rings 142 and 144and to prevent undesired contact between the turbine blades and theinner radial faces of constraint rings 142 and 144. In one embodiment,spacer members 146 are each fabricated from the same or similar material(e.g., a nickel-based superalloy) as are the bladed pieces that make-upformation 132. The CTE of spacer members 146 will also typically begreater than the CTE of constraint ringers 142 and 144.

As shown in FIGS. 7 and 9, a first sacrificial crush ring 160 may bedisposed around a first circumference of ring formation 132 and betweentooling contact surface 126 (identified in FIG. 9) and thermal growthconstraint ring 142; and a second sacrificial crush ring 162 may bedisposed around a second circumference of ring formation 132 and betweentooling contact surface 128 (identified in FIG. 9) and thermal growthconstraint ring 162. Crush rings 160 and 162 are fabricated from amaterial that softens and plastically deforms during thermal processing.In this manner, the material from which crush rings 160 and 162 areformed can flow or deform to accommodate misalignment between the bladedpieces and geometric imperfections in tooling contact surfaces 126 and128 and thereby ensure that a substantially uniform load is applied toring formation 132 during thermal bonding. In addition to softening atthermal processing temperatures, the material from which crush rings 160and 162 are formed also preferably has a CTE similar to the materialfrom which the bladed pieces are fabricated to allow crush rings 160 and162 to grow radially outward in conjunction with ring formation 132.Finally, it is also desirable for the crush ring material to berelatively low cost in view of the consumable or sacrificial nature ofcrush rings 160 and 162. In one implementation, crush rings 144 are eachfabricated from a low carbon or other low superalloy steel, such as anAISI 1010 or 1020 steel.

FIGS. 10-13 generically illustrate the manner in which thermal growthconstraint tool 140, and specifically thermal growth constraint rings142 and 144, impede the outward radial growth of ring formation 132during the directed thermal growth bonding process. While only thermalgrowth constraint ring 142 is shown in FIGS. 10-13 and described below,it will be appreciated that the following description applies equally tothermal growth constraint ring 144. With reference to FIG. 11, which isa detailed view of the region of ring formation 132 encompassed bydashed circle 163 (FIG. 10), the outer circumferential surface of crushring 160 may be separated from the inner circumferential surface ofconstraint ring 142 by a small radial gap 164 prior to heating. As ringformation 132 and thermal growth constraint tool 140 are heated to thedesired bonding temperature, ring formation 132 grows in a radiallyoutward direction such that the outer circumference of formation 132gradually increases. In embodiments wherein crush ring 160 is formedfrom a material having a CTE similar or identical to that of ringformation 132, crush ring 160 likewise grows radially outward inconjunction with ring formation 132. Constraint ring 142 grows radiallyoutward more gradually than does ring formation 132 and crush ring 160;consequently, radial gap 164 eventually closes and crush ring 160 comesinto to contact with constraint ring 142 (shown in FIG. 13). At thispoint, further outward radial growth of ring formation 132 and crushring 160 is impeded or physically obstructed by constraint ring 142.Further growth of ring formation 132 consequently results in theapplication of a significant convergent pressure urging shanks 118together, as represented in FIG. 13 by arrows 166. A strongmetallurgical bond is formed between neighboring shanks 118 as a resultof this combination of high temperatures and high convergent pressures.The bond between neighboring shanks 118 may be created by melting (andthe subsequent re-solidification) of a braze material disposed betweenshanks 118, if provided, by diffusion bonding, or by a combinationthereof.

With continued reference to FIGS. 11-13, in embodiments wherein theshanks of neighboring bladed pieces 116 are to be joined by brazing, astrip or strip of braze foil 168 may be disposed between neighboringshank bonding surfaces. In this event, the bonding temperature ortemperatures to which ring formation 132 and thermal growth constrainttool 140 are heated during STEP 138 (FIG. 4) will typically be selectedto exceed the melt point of braze superalloy material, while being lessthan the melt point of the other components included within tooling 118and ring formation 132. Conversely, in embodiments wherein a braze foilor other bonding material is not disposed between neighboring shanks,the bonding temperatures and components dimensions are selected tocreate sufficient convergent pressures to ensure the creation of adiffusion bond between neighboring shanks. In many cases, both a brazejoint and a diffusion bond may be created between neighboring shanksduring the above-described directed thermal growth bonding process. Byway of non-limiting example, ring formation 132 and thermal growthconstraint tool 140 may be heated to temperature approaching orexceeding about 2150° Fahrenheit during the above-described bondingprocess.

Through the above-described directed thermal growth process, bladedpieces 116 are metallurgically consolidated into a monolithic bladedring, which, in this case, serves as forward bladed ring 94 (FIG. 3) ofaxially-split radial turbine 20. Next, at STEP 170 of exemplaryfabrication method 110 (FIG. 4), forward bladed ring 94 is bonded onto aforward rotor disk to produce forward turbine rotor 50 (FIGS. 1-3). Anexample of a forward rotor disk 172 onto which forwarded bladed ring 94may be bonded is shown in FIG. 14. Forward rotor disk 172 may beproduced by, for example, forging a cylindrical preform from a selectedsuperalloy or other material and subsequent machining the forging to anear-net shape. Forward rotor disk 172 is preferably fabricated from ansuperalloy tailored to the unique operational environment of the rotordisk; e.g., an superalloy having relatively high stress rupture and lifecycle fatigue properties. In one implementation, forward rotor disk 172is fabricated from an oxidation-resistant, nickel-based superalloy;while the forward bladed ring is fabricated from a single crystalsuperalloy of the type described above.

After production of the rotor disk forging, the outer annular portion offorward rotor disk 172 may be machined to impart disk 172 with a desiredsurface finish and an outer diameter approximately equivalent to theinner diameter of the machined airfoil ring. The leading and trailingannular interfaces between the outer circumference of the ring and theinner circumference of disk 172 are sealed by, for example, brazing toprevent the ingress of air into the ring-disk interface that mayotherwise interfere with optimal bonding. A hot isostatic pressing (HIP)diffusion process or other bonding process can then carried-out to bondforward bladed ring to forward rotor disk 172 and thereby yield apartially-completed version of forward turbine rotor 50, as shown inFIG. 15. Additional machining steps may be performed to completefabrication of the forward turbine rotor; e.g., the formation ofinter-blade cut-outs or other strain relief features and the removal ofthe braze material, if utilized to inter-bond the bladed pieces (STEP176, FIG. 4). After production, forward turbine rotor 50 may beinspected for structural defects utilizing any one of a number of knownnon-destructive inspection techniques.

Next, at STEP 178 (FIG. 4), aft turbine rotor 52 (FIGS. 2 and 3) isproduced. Aft turbine rotor 52 can be produced utilizing essentially thesame steps employed in the fabrication of forward turbine rotor 50. Forexample, and referring once again to FIGS. 2 and 3, a number of aftbladed pieces, which include aft blade segments 62, may be individuallycast from a chosen superalloy (e.g., a single crystal superalloy of thetype described above) and subsequently consolidated into an aft bladedring 100, which is then bonded onto a separately-produced aft rotor disk102. In this case, aft bladed ring 100 may be fabricated from the samematerial (e.g., a single crystal superalloy) as is forward bladed ring94 or, instead, fabricated from a different material. Aft rotor disk 102may likewise be fabricated from the same material as is forward rotordisk 96 utilizing, for example, an isothermal forging process. Finally,as indicated in FIG. 4 at STEP 180, the forward and aft turbine rotorsare assembled to produce axially-split radial turbine 20 (FIGS. 1-3),which can then be installed within a gas turbine engine; e.g., asdescribed above, a curvic, tie-shaft, or other coupling may be utilizedto axially-clamp forward turbine rotor 50 to aft turbine rotor 52 andthereby yield the assembled axially-split radial turbine 20.

It should thus be appreciated that, by way of the above-describedfabrication process, forward bladed ring 94 (and therefore the entiretyor at least a substantial portion of each forward blade segment 58 shownin FIGS. 1-3) can be fabricated from a first material (e.g., a singlecrystal superalloy) highly resistant to oxidation and other forms ofchemical degradation, while forward disk 96 is fabricated from a secondmaterial (e.g., an equiaxed superalloy) having high stress rupture andlow cycle fatigue properties. Similarly, aft bladed ring 100 (andtherefore aft blade segments 62) can be fabricated from a thirdsuperalloy (e.g., a single crystal superalloy), while forward disk 96 isfabricated from a fourth material specifically tailored to its operatingconditions. In a preferred embodiment, forward bladed ring 94 and aftbladed ring 100 are each fabricated from the same or similar singlecrystal alloys; while forward disk 96 and aft disk 102 are fabricatedfrom the same or similar non-single crystal alloys. In still furtherembodiments, aft turbine rotor 52 may consist of a single material andmay be fabricated as a unitary casting or forging.

The foregoing has thus provided embodiments of an axially-split radialturbine suitable for usage in a gas turbine engine that is relativelylightweight and cost effective to implement. In the above-describedembodiments, the axially-split radial turbine included internal coolingfeatures (e.g., blade cooling passages) favorably increasing the thermalcapabilities of the radial turbine. The foregoing has also providedembodiments of a method for fabricating such a radial turbine enablingdifferent turbine sections to be fabricated from disparate alloystailored to the particular operating conditions experienced by eachturbine section. Notably, embodiments of the above-described method forproducing a radial wheel utilize multiple separately-cast piecesfacilitate individual casting of blades having integral coolingpassages. In so doing, axially-split radial turbines produced inaccordance with the above described fabrication methods improve highcasting yield and enable the separate pieces to be produced fromsuperalloy compositions tailored to the varying thermal-mechanicalrequirements of a wide variety of applications.

While at least one exemplary embodiment has been presented in theforegoing Detailed Description, it should be appreciated that a vastnumber of variations exist. It should also be appreciated that theexemplary embodiment or exemplary embodiments are only examples, and arenot intended to limit the scope, applicability, or configuration of theinvention in any way. Rather, the foregoing Detailed Description willprovide those skilled in the art with a convenient road map forimplementing an exemplary embodiment of the invention. It is beingunderstood that various changes may be made in the function andarrangement of elements described in an exemplary embodiment withoutdeparting from the scope of the invention as set-forth in the appendedclaims.

What is claimed is:
 1. An axially-split radial turbine, comprising: a forward turbine rotor, comprising: a forward bladed ring; and a forward disk joined to the forward bladed ring; and an aft turbine rotor; a plurality of hub flow paths extending from the forward turbine rotor to the aft turbine rotor, the plurality of hub low paths each comprising: a leading portion defined, at least in substantial part, by the forward disk; and an intermediate portion defined, at least in substantial part, by the forward bladed ring; wherein the forward turbine rotor and the aft turbine rotor are disposed in an axially-abutting, rotationally-fixed relationship.
 2. The axially-split radial turbine of claim 1 wherein the forward bladed ring comprises a plurality of forward bladed pieces, and wherein the plurality of forward bladed pieces are cast to include a plurality of blade cooling passages.
 3. The axially-split radial turbine of claim 2 wherein the aft turbine rotor comprises an aft disk; and wherein the axially-split radial turbine rotor further comprises an inner disk cavity formed between the forward disk and the aft disk, the inner disk cavity in fluid communication with the plurality of blade cooling passages, the inner disk cavity at least partially bound by the forward bladed ring, the forward disk, and the aft turbine rotor.
 4. The axially-split radial turbine of claim 3 further comprising at least one hub cooling passage extending from an inner circumferential surface of the axially-split radial turbine to the inner disk cavity.
 5. The axially-split radial turbine of claim 1 wherein the forward bladed ring is produced from a plurality of forward bladed pieces, the plurality of forward bladed pieces individually cast and metallurgically consolidated utilizing a direct thermal growth process to produce the forward bladed ring.
 6. The axially-split radial turbine of claim 5 wherein the plurality of forward bladed pieces are each cast from a single crystal superalloy, and wherein the forward disk is produced from a non-single crystal superalloy.
 7. An axially-split radial turbine comprising: a forward turbine rotor, comprising: a forward disk; and a forward bladed ring circumscribing and metallurgically bonded to the forward disk at a bond line that terminates adjacent axially opposed end portions of the forward turbine rotor, as taken along a rotational axis of the axially-split radial turbine; and an aft turbine rotor disposed axially adjacent to and rotationally fixed relative to the forward turbine rotor; wherein the forward bladed ring comprises a plurality of forward bladed pieces, and wherein each of the plurality of forward bladed pieces is individually cast to include at least one forward blade segment.
 8. The axially-split radial turbine of claim 7 further comprising: a plurality of blade cooling passages extending in the plurality of forward bladed pieces; and an inner disk cavity formed within the axially-split radial turbine and in fluid communication with the plurality of blade cooling passages.
 9. The axially-split radial turbine of claim 7 wherein the plurality of forward bladed pieces are individually cast from a single crystal alloy, and wherein the forward disk is composed of a non-single crystal alloy.
 10. The axially-split radial turbine of claim 7 wherein the aft turbine rotor comprises: an aft bladed ring produced from a plurality of aft bladed pieces each including at least one aft blade segment; and an aft disk joined to the aft bladed ring.
 11. The axially-split radial turbine of claim 10 further comprising: a plurality of blade cooling passages extending in the plurality of aft bladed pieces; and an inner disk cavity formed within the axially-split radial turbine and in fluid communication with the plurality of blade cooling passages.
 12. The axially-split radial turbine of claim 7 wherein the bond line has a generally conical geometry, which decreases in outer diameter when moving in a fore-aft direction.
 13. The axially-split radial turbine of claim 7 further comprising a hub flow path extending over the aft and forward turbine rotors, the hub flow path having a leading portion defined by the forward disk and having an intermediate portion defined by the forward bladed ring.
 14. An axially-split radial turbine, comprising: an aft turbine rotor, comprising: an aft disk; and an aft bladed ring circumscribing and metallurgically bonded to the aft disk at a bond line that terminates adjacent axially opposed end portions of the aft turbine rotor, as taken along a rotational axis of the axially-split radial turbine; and a forward turbine rotor disposed axially adjacent to and rotationally fixed relative to the aft turbine rotor; wherein the aft bladed ring comprises a plurality of aft bladed pieces, and wherein each of the plurality of aft bladed pieces is individually cast to include at least one aft blade segment.
 15. The axially-split radial turbine of claim 14 wherein the plurality of aft bladed pieces are individually cast from a single crystal alloy, and wherein the aft disk is composed of a non-single crystal alloy.
 16. The axially-split radial turbine of claim 14 wherein the forward turbine rotor comprises: a forward bladed ring produced from a plurality of forward bladed pieces each including at least one forward blade segment; and a forward disk joined to the forward bladed ring.
 17. The axially-split radial turbine of claim 16 wherein the plurality of forward bladed pieces include a plurality of blade cooling passages.
 18. The axially-split radial turbine of claim 17 further comprising an inner disk cavity formed between the forward disk and the aft disk, the inner disk cavity in fluid communication with the plurality of blade cooling passages. 